Conference PaperPDF Available

Hawai'iSat-1: Development Of A University Microsatellite For Testing a Thermal Hyperspectral Imager

August 2010

DOI:10.2514/6.2010-8922

Conference: AIAA SPACE 2010 Conference & Exposition

Authors:

Trevor Sorensen

University of Hawaiʻi at Mānoa

Lloyd French

University of Hawai'i System

Jeremy K Chan

University of Hawai'i System

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HawaiʻiSat-1 Project Organization

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HawaiʻiSat-1 Project Schedule

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HawaiʻiSat-1 System Context Diagram

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EPS Block Diagram

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Battery Charge Status Over One Week

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Content uploaded by Miguel A Nunes

Content may be subject to copyright.

American Institute of Aeronautics and Astronautics

HawaiʻiSat-1: Development Of A University Microsatellite

For Testing a Thermal Hyperspectral Imager

Trevor C. Sorensen, Lloyd French, Jeremy K. Chan, William K. Doi, Elizabeth D. Gregory, Marcelo H. Kobyashi,

Zachary K. Lee-Ho, Miguel Nunes, Eric J. Pilger, Reid A. Yamura, Lance K. Yoneshige

Hawaiʻi Space Flight Laboratory, University of Hawaiʻi, Honolulu, HI, 96822

The Hawaiʻi Space Flight Laboratory (HSFL) was established at the University of

Hawaiʻi (UH) at Manoa for two primary purposes: (1) to educate students and help prepare

them to enter the technical workforce, and (2) to help establish a viable space industry that

will benefit the State of Hawaiʻi. The second HSFL space mission, currently scheduled to be

launched in late 2012, is STU-2, which includes a spacecraft being designed and built by the

HSFL called HawaiʻiSat-1. The Operationally Responsive Space (ORS) Office located at

Kirtland Air Force Base in New Mexico oversees the LEONIDAS contract, under which the

STU-2 mission and the HawaiʻiSat-1 satellite are being developed.

The primary objectives for HawaiʻiSat-1 mission are: (1) to demonstrate the ability of the

HSFL to design, build, and operate a small satellite in the 80-kg class as a platform to test

new technologies; (2) support the C-band Radar Transponder Experiment (CRATEX)

Payload; (3) support the testing of the Thermal Hyperspectral Imager being developed at

UH; and (4) perform Earth imaging using the HSFL Imaging Payload. The CRATEX

payload, provided by Vandenberg Air Force Base, uses C-band transponders and precise

orbit determination (provided by onboard GPS receivers) to help the Department of Defense

and NASA calibrate their C-band tracking radars around the world. The HawaiʻiSat-1

spacecraft will be placed into a 550-km circular 9 p.m. ascending Sun Synchronous Orbit to

optimize its support of the CRATEX payload.

The 85-kg HawaiʻiSat-1 spacecraft is 3-axis stabilized using three magnetic torque rods

and a reaction wheel for attitude control; and three sun sensors plus two inertial

measurement units (each including a 3-axis magnetometer) for attitude determination.

Communication is provided by S-band and UHF-band transceivers linked to a ground

station located in the Kauaʻi Community College in Hawaiʻi, and other partner ground

stations. Control of the mission will be done in the HSFL Mission Operations Center located

on the University of Hawaiʻi campus at Manoa. Integration and testing of the spacecraft will

be done in the clean rooms at the HSFL facilities on the UH campus, which includes a 1.6

meter diameter thermal vacuum chamber.

The HSFL is using a core team of experienced professionals supplemented with graduate

and undergraduate students to design, build, and test the HawaiʻiSat-1 spacecraft within a

period of approximately two years from System Requirements Review until ready for

launch. To keep the spacecraft cost to a minimum, commercial-off-the-shelf (COTS)

components will be used when possible. The fairly benign radiation environment of such a

low altitude orbit and the use of aluminum sheeting to shield the critical avionics, make the

risk of using COTS for a 2-3 year mission to be acceptable while greatly reducing the cost as

compared to using space-hardened parts.

I. Introduction

he Hawaiʻi Space Flight Laboratory (HSFL) was established at the University of Hawaiʻi at Manoa (UH) for

two primary purposes: (1) to educate students and help prepare them to enter the technical workforce, and (2) to

help establish a viable space industry that will benefit the State of Hawaiʻi.

The Low Earth Orbit Nanosat Integrated Defense Autonomous Systems (LEONIDAS) Program is a

Congressionally sponsored program which is of importance to both the United States and to the State of Hawaiʻi. Its

purpose is to develop and demonstrate small-satellite orbital launch capability from the Pacific Missile Range

Facility (PMRF). The objectives of the program include establishing a technical work force in the State of Hawaiʻi

with a development program that trains students to enter the technical work force. The original vision for the

program was to develop a chain of small satellites to pass information to a single ground station and assist with

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Hawai‘i Space Flight Laboratory

Management

and Facilities

Mission Oversight

PM & SE

System I&T

S/C -> PAD I&T

Launch Support

Prime Ground

Segment

Ground Station

Operations Center

Payload Customers

CRATEX (VAFB)

HIP (HSFL)

THI (HIGP)

SPARK Launch Services

SPARK I

NASA ARC &

Santa Clara University

PM & SE

Design and

Fabrication

SPARK PAD

SC EPS

SC Telecom

SC ADCS

B/U MOC+GS

Sponsors/Support

PM/SE/GS Support

EPS,ADCS,Telecom

Hawai‘iSat-1 Satellite

Support

Launch

Support &

Payload H/W

Hawai‘iSat-1 Mission Architecture

Mission

Services

System Test

& Integration

PL Data

Future GS Site

Secondary Ground

Segment

Design & Fab.

EPS, COMM, TCS,

OBCS, FSW, S&M

Subsystem

HW/SW

Ground Station Services

SPARK PAD

Support

Figure 1. STU-2/HawaiʻSat-1 Mission Architecture

disaster-relief efforts. It later evolved into a constellation of very small (nano-) autonomous satellites to aid the

Department of Defense (DoD). By the time funding was approved in 2007, the program was designated to cover two

launches and two spacecraft. Oversight of this program was given to the U.S. Army’s Space and Missile Defense

Command (SMDC) with headquarters in Huntsville, AL. In February 2008 the oversight of the program was

transferred to the Operationally Responsive Space (ORS) Office with headquarters at the Kirtland Air Force Base in

Albuquerque, NM.

The first LEONIDAS mission was to consist of a small (30 – 50 kg) satellite to demonstrate a stable platform for

remote sensing (especially missile detection and tracking). The LEONIDAS Missions #1is designated as Science

and Technology for the University (STU) #1. This is the official mission launch designation provided to PMRF. The

first launch of the LEONIDAS Program and HSFL is STU-1, which included the LEO-1 spacecraft; other secondary

satellites (CubeSats); the Super Strypi launch vehicle, being developed by Sandia National Laboratories for HSFL;

and the payload adapter and deployer (PAD), designed and built by the NASA Ames Research Center.

With the transfer of oversight of the LEONIDAS Program from SMDC to ORS, the mission for the first satellite,

LEO-1, also changed. Originally it was designed to show the suitability of the microsat bus to be a platform for

remote sensing of missile launches and tracking. It then was changed to be a microsat platform to test various

technologies of interest to the Department of Defense including the ORS office. During the six months following the

switchover, new experimental payloads were found and feasibility studies were performed. Finally the payload

manifest was finalized with the inclusion of the C-band radar transponder experiment (CRATEX) for the USAF; the

Coherent Electromagnetic Radio Tomography (CERTO) experiment for the Naval Research Laboratory (NRL); and

the LEO-1 Solar Array Experiment (LSAE). The LEO-1 design successfully passed its Preliminary Design Review

(PDR) in May 2009. The LEO-1 satellite was reported in a paper presented at AIAA Space 2009 Conference.1

However, in late 2009 the LEO-1 satellite project was canceled due to funding concerns, and the first two satellite

missions, LEO-1 and LEO-2 were combined into a single satellite mission, called HawaiʻiSat-1 and moved to the

second launch of the new SPARK launch vehicle under development by HSFL. With the combining of the missions,

there were some changes in the payloads, with two LEO-1 experiments (CERTO and LSAE) being dropped while

the major payload from LEO-2 (thermal hyperspectral imager) was added. HawaiʻiSat-1 successfully completed its

PDR in June, 2010 and is scheduled for completion and ready for launch at the end of 2011. This paper presents the

baseline design of HawaiʻiSat-1 after its PDR.

The HawaiʻiSat-1 is a three-axis stabilized octagonal spacecraft of approximately 85 kg mass and will be

launched into a 550-km altitude 9 p.m. ascending node sun synchronous orbit (SSO). Communication is provided by

a UHF- transceiver linked to a ground station located in the Kauaʻi Community College (KCC) on the Hawaiʻian

island of Kauaʻi and other partner ground stations around the world. Control of the mission will be done in the

HSFL Mission Operations Center located on the University of Hawaiʻi campus at Manoa.

II. Mission Description

This section describes the mission architecture, project management, the launch trajectory, and orbits analysis.

A. Mission Architecture

The HawaiʻiSat-1

spacecraft is part of the

designated STU-2 launch

from PMRF. Figure 1 shows

the top level mission

architecture for STU-2.

HSFL, located at the UH

Manoa campus, is at the

center of the system and will

be responsible for design,

fabrication, project

management; integration and

testing of both the

HawaiʻiSat-1 spacecraft and

the PAD. All the payloads

being flown on STU-2 will

be integrated on the PAD at

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HSFL before shipment to PMRF for launch vehicle integration. HSFL will also be responsible for the ground

segment including the ground stations, and mission operations centered in the HSFL Mission Operations Center

(HMOC).

B. Project Management

1. Organization and Staffing

The HSFL employs a small and dedicated team to lead the activity of developing the HawaiʻiSat-1 spacecraft.

To fulfill the university’s mission to educate students, HSFL involves several graduate students for these projects.

The HSFL management has thus hired a core team of experienced experts in the various fields required to complete

the development effort. This cadre of experts serves as the mentors for the students and helps mitigate the risks

inherent to the development of a new launch vehicle and spacecraft. In an effort to manage developmental risk on a

multi-million dollar highly technical project such as HawaiʻiSat-1, HSFL leverages talented students from CubeSat

project programs through Hawaiʻi Space Grant Consortium and the Native Hawaiʻian Science and Engineering

Mentoring Program.

The HSFL is an element of the University of Hawaiʻi and thus has a primary purpose to educate st udents through

workforce development. This purpose is reflected in the organizations employed for the HawaiʻiSat -1 project. By

employing a number of undergraduate and graduate students, many of whom that have matriculated through

CubeSat team training programs, the HSFL will accomplish the education of students to prepare them to enter the

workforce as already experienced and productive engineers and scientists, and to help keep the development costs of

the projects low. The project management and key subsystem lead engineers are all experienced full-time HSFL

staff or part-time faculty. The experience of the key personnel allows them to perform multiple tasks within the

organization with the help of undergraduate and graduate students. In some cases the lead engineers will do the work

themselves, usually where a high level of technical expertise or experience is required, or they will supervise the

performance of tasks by students. Even when the lead engineers do the task themselves, one or more students will be

involved if at all possible to observe and learn. The staffing level for most of the project, including staff engineers,

faculty, and students, stays in the range of 8 to 9 Full Time Equivalents (FTEs).

Project managers, Lloyd French and Dr. Trevor Sorensen lead the HawaiʻiSat-1 project organization and both

report to the HSFL Director, Dr. Luke Flynn. The organization of the HawaiʻiSat-1 project is shown in Figure 2.

Figure 2. HawaiʻiSat-1 Project Organization

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The HSFL project managers and HSFL project systems engineer are collectively considered the Project

Management Team. The Project Management Team will lead the Spacecraft Team and Mission Support Team

(MST) through all the phases of development towards delivery. The HSFL Spacecraft Team consists of spacecraft

subsystem engineers and payload engineers. The Mission Support Team is composed of HSFL facilities managers

in integration and testing, ground station, mission operations, and launch support. The project manager and project

system engineer manage budgets, schedules, and tasks for the project. HSFL also provides service support for the

project with administration, contracts, and data management and archiving. HSFL has instituted a Change Control

Board (CCB) to help manage changes to the design, risk management, mission assurance, documentation, and

configuration control for documentation and software. These methodologies are based on the standard processes of

NASA and the Department of Defense, but are generally simplified as appropriate for a small project.

The project management team will plan for project integration and test (I&T) activities through the I&T

Manager. The I&T Manager serves two functions: (1) Project I&T, and (2) Facilities I&T. The HSFL I&T

Manager works with the project and provides integration and testing support. To facilitate training and workforce

development, the spacecraft engineers also perform environmental tests and functional tests for hardware. The

spacecraft team will work as the I&T team to complete fabrication, assembly, integration, and testing. The project

management team plans for mission operations coordinating with the Mission Operations Manager. The

HawaiʻiSat-1 mission will have operational support of command, telemetry, and data management. The Ground

Segment Manager coordinates with the Mission Operations Manager. Ground stations are directed by the project

management team to support HawaiʻiSat-1 ground telecommunication needs for command and telemetry with data

interfaces with mission operations. The Launch Support Manager, who works with the project management team,

will plan HawaiʻiSat-1 support with the launch vehicle provider and launch range.

2. Development Process

The basic design process being followed by the HSFL Team is fairly standard for similar space projects. It is

based loosely on the design process outlined in Reference 2. The development process has been divided into five

phases as defined in Table 1. The HSFL phase nomenclature is based on the NASA system (A-E) while the

corresponding DoD nomenclature is in parentheses (I-V). Several milestones have been planned into the

development process including the following:

1. System Requirements Review (SRR)

2. Preliminary Design Review (PDR)

3. Critical Design Review (CDR)

4. Test Readiness Review (TRR)

5. Mission Readiness Review (MRR)

6. Pre-Ship Review (PSR)

Table 1. HawaiʻiSat-1 Development Process Phases

Phase Name Description

A (0) Concept Definition Defines requirements and baseline conceptual design – ends in SRR

B (I) Definition & Improves on baseline design with further analyses and trades and development

Acquisition Planning of prototype hardware – ends in PDR with preliminary design

C (II) Detailed Design Completion of system design and development FlatSat hardware – ends in CDR

D (III) Development Includes procurement, fabrication, integration & testing of flight hardware – ends in

PSR to launch

E (III) Operations All operations after launch including Launch & Early Orbit (L&EO), Engineering

Evaluations & Checkout (EE&C), nominal operations, and terminal operations – ends

with reentry or loss of satellite

The developmental process for Phases A & B being followed by HSFL in the development of HawaiʻiSat-1

utilizes unique team dynamics through concurrent engineering and forward hardware prototyping. The concurrent

engineering methodology is a key activity to aid in design data fusion with multiple subsystems. The design issues

are worked on in real time, thus the turn around time to implementation is shortened. This method is very useful on

projects that are cost constrained. The shortened schedule reduces labor costs.

The HawaiʻiSat-1 project schedule is shown in Figure 3. The project uses a forward hardware prototyping

methodology to reduce hardware development risk and control cost by shortening mission development time. The

spacecraft team prototypes key subsystems deemed risky during Phase B thus aiding the team’s understanding of

spacecraft design’s robustness. By having prototypes in Phase B, the HawaiʻiSat-1 development creates hardware

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parallel efforts with design. Advanced understanding of the hardware carries over to the design effort. Thus the

design and hardware become complementing parallel efforts to reduce risk. The spacecraft is currently scheduled to

be completed by December 2011.

2. Systems Engineering

The role of systems engineering in the project is to define and develop an operable system capable of meeting

mission requirements within the imposed constraints, including performance, cost, risk, and schedule. The HSFL

approach follows common systems engineering best practices appropriate to the project and include the following

phases: functional analysis; system synthesis; system evaluation and decision; and system definition. The Systems

Engineering Management Plan document captures in greater detail the approach to the systems engineering effort.

The project systems engineering effort also includes configuration management; oversight of project

documentation; and places a major role in the project risk management strategy. To group various activities and

system products, Figure 4 shows the baseline HawaiʻiSat-1 mission systems context. The system is divided into

four major segments: launch, ground, satellite, and customer. Each segment has a major role in HawaiʻiSat-1

mission operations.

The satellite segment includes both the satellite bus, and its payloads. The bus is designed to be a general

purpose multi-mission bus. The goal is to gain flight heritage on this bus, and be able to accommodate future

missions. More information both the bus and payload can be found in following satellite segment sections.

The launch segment is mostly handled by the launch services provided by the STU-2 mission. Coordination is

performed to accommodate for range safety requirements, interface control, and pre-launch checkout procedures.

STU-2 launch services will provide the equipment and personnel who will ultimately carry out the satellite’s launch.

The ground segment encompasses all ground stations (GS) and the HSFL mission operations center (HMOC) in

Hawaiʻi. The ground stations include satellite commanding and telemetry receiver variants. Partnerships are

currently being explored to expand the ground segment to provide more coverage and robustness. In addition, the

ground segment contains the mission operations center which coordinates all the segment ground stations. Since the

Internet connection between the ground stations and HMOC may not reliable, the ground stations are designed to

operate independently if necessary. When an Internet connection is available, the HMOC can synchronize ground

segment activities. The HMOC will also be responsible for sparse downlinked payload data reassembly (Level 0

processing). With the payload data reconstituted on the ground, the HMOC will then be able to deliver data

products to the customer segment. The HMOC also works with the customer segment for payload operations

scheduling.

Figure 3. HawaiʻiSat-1 Project Schedule

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Satellite Bus

On-Orbit Automation

Communications

Tlm/Cmd Processing

Satellite Payloads

CRATEX

THI

HIP

Satellite Segment

Ground Stations

Backup SC C&C

Satellite Tracking

Telemetry Services

Telecommand Services

Remote Operation

Encrypted MOC Conn.

Ground Segment

Mission

Operations Center

Mission Planning

Contact Execution

Data Management

Analysis

SC C&C

Launch Vehicle Services

Logistics

SPARK LV Integration

Pre-flight Checkout

Launch Monitoring

Orbit Determination

Range Safety

Launch Segment

Payload Customers

VAFB (30th SW)

HIGP

HSFL

Customer Segment

Telecommands

(UHF):

- Automation

Schedule

- Real Time

Telemetry (S-Band,UHF):

- Real Time

- Buffered Payload Data

Ground Station Commanding

Satellite Commanding

Satellite Telemetry

Ground Station Telemetry

(Encrypted Internet Link)

Schedule

Requests

(via Encrypted

Internet Link)

Payload Data

from Satellite

(via Encrypted

Internet Link)

One-Shot Transfer:

- Initial Orbit Data

- Launch Footage & Data

C-Band Transponding (CRATEX)

HawaiiSat-1 System Context Diagram

Figure 4. HawaiʻiSat-1 System Context Diagram

C. Trajectories and Orbit Analysis

1. Launch Profile

The HawaiʻiSat-1 spacecraft will be ready for launch by December 2011. It will be launched by a SPARK

(enhanced Super Strypi) three-stage, spin-stabilized, solid-propellant rocket from the Pacific Missile Range Facility

(PMRF) located on Kauaʻi in a southerly retrograde direction. The notional launch profile is shown in Figure 5.

The spacecraft will be deployed from the Payload Adapter and Deployer (PAD), being developed by NASA

Ames Research Center. The PAD will be capable of carrying one primary microsatellite and up to 24 or 32 cubesats

or equivalents. The nominal insertion orbit is into a 9 p.m. ascending circular sun synchronous orbit (SSO) with an

altitude of 550 km. and inclination of ~ 97.6º. After the third stage combustion has terminated, a yo-yo mechanism

will despin the vehicle from 1 rps to no more than 4º per second.

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2. Orbit Analysis

The HawaiʻiSat-1 spacecraft has no propulsion subsystem and will be unable to alter its orbit. It is expected to

reenter the Earth’s atmosphere within 25 years after the end of the design mission (three years). The orbital analysis

was performed at HSFL based upon orbit insertion data provided by Sandia National Laboratories. The assumed

baseline orbit is shown in Table 2. The design length of the mission is three years. Dispersion analyses were also

done, but are not included in this paper.

Three ground stations were assumed for the initial study: Kauaʻi HI, Santa Clara CA, and Fairbanks AK.

Assuming an elevation of 10 degrees for the line of sight the total contact duration during a week will be: 1.4 hr/wk

Launch Date 1 January 2012 19:55:40 UTC

Altitude 550 km

Orbit Type Circular, Sun Synchronous

Inclination 97.5976 degrees

Circular Velocity 7.585 km/s

Orbit Angular Velocity 3.723 deg/min

Period 95.6 minutes

Revolution per Day 15.06

Force Model Parameters 21 x 21 EGM96 Earth gravitation

Solar and lunar gravitation

CD = 2.2

A/m = 0.005 m2/kg

Jacchia 1970 Lifetime atmosphere

PAD

(8 P-Pods)

~ 8g expected

PAD

(8 P-Pods)

~ 8g expected

Figure 5. Notional Enhanced SPARK Launch Profile3

Table 2. Baseline Parameters for Orbital Analysis

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for the ground station in Kauaʻi, 1.6 hr/wk for Santa Clara and 4.2 hr/wk for Fairbanks. Figure 6 shows the nominal

orbit ground traces of the first three complete orbits of the HawaiʻiSat-1. The importance of a ground station in

Alaska is evident from these results. The altitude prediction is shown in Figure 7, the inclination prediction in Figure

8, and the descending node solar time (local time) over the three-year lifetime is shown in Figure 9. This latter plot

is important for a desired Sun Synchronous Orbit (SSO) to show how well the satellite keeps to the desired local

time. Figure 10 shows how long after launch the satellite comes into contact with the three ground stations,

indicating a fairly large gap before we first see the satellite after deployment. HSFL is attempting to arrange for a

ground station in Europe to fill this gap. The HawaiʻiSat-1 orbits will have approximately one third of the orbit in

eclipse (umbra) as shown in Figure 11. The lifetime prediction analysis (Figure 12) predicts the decay for this

satellite in December of 2035 which is about 24 years of orbital lifetime from the launch date. This shows that active

deorbiting is not needed.

Figure 7. Orbital Altitude Prediction

Figure 6. Ground Traces of First Three Nominal Orbits

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Figure 8. Orbital Inclination Prediction

Figure 9. Descending Node Solar (Local)Time

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Figure 10. Early Ground Station Contacts

Figure 11. Percentage of Orbit in Eclipse

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III. Spacecraft Description

The HawaiʻiSat-1 is a multi-mission microsatellite which is being designed for low earth orbit operation.

Figures 13 and 14 show the preliminary satellite packaging.

Rear Nadir View

-Y

-X

Sunsensors

(3x)

C-band

Antennas (2x)

GPS Antennas

(2x)

Telecom

Antennas

(3x)

15”

Lightband

Separation

Connector

Diagnostic

Port

Front Zenith View

SIP

Camera

HIP

Camera

THI

Cutout

Rear Nadir View

-Y

-X

Sunsensors

(3x)

C-band

Antennas (2x)

GPS Antennas

(2x)

Telecom

Antennas

(3x)

15”

Lightband

Separation

Connector

Diagnostic

Port

Front Zenith View

SIP

Camera

HIP

Camera

THI

Cutout

Figure 13. HawaiʻiSat-1 External View

Altitude

Eccentricity

Launch L+24 yr

Altitude

Eccentricity

Altitude

Eccentricity

Launch L+24 yr

Figure 12. Orbital Lifetime Plot

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The satellite features body mounted solar panels on nine of its ten sides. GPS antennas are mounted on the

nominally zenith side of the spacecraft (-Z) for attitude determination and payload operations. Concept antennas are

placed on the satellite, but will be finalized when more detailed radiation pattern analysis is performed. Two

directional C-band antennas are mounted on the nominally nadir side of the spacecraft for the C-band transponders.

Several holes are cut in the nadir (+Z) surface to accommodate for the imagers.

The satellite is designed to be the primary payload aboard HSFL SPARK series launch vehicles. Thus, a 15”

upper lightband ring is mounted on the nadir surface for primary payload satellite deployment. As the bus becomes

more mature, the 15” lightband ring will also allow for mounting and deployment from an ESPA ring secondary

payload interface. Since horizontal integration onto a lightband ring receptacle is expected, the entire spacecraft is

designed to be handled in any orientation.

B. Bus Subsystems

1. System

Figure 15 shows the spacecraft bus interconnections. There are three primary subsystems which run at all times:

(1) The on-board computer subsystem (OBCS) is the main flight computer which runs the flight software (FSW).

The OBCS interfaces with all other bus subsystems using cost effective COTS interfaces including Ethernet, RS232,

and an array of analog interfaces specifically for RTD temperature sensors. (2) The EPS subsystem is at the center

of the spacecraft power generation, storage, control, and distribution. A single point ground is implemented via the

EPS subsystem. (3) Finally, the telecom subsystem handles all satellite to ground segment communications.

Secondary subsystems include the attitude determination and control subsystem (ADCS), which provides three

axis stabilization. The ADCS subsystem is using a new control and estimation algorithm which will be proven on

the satellite’s maiden voyage. The thermal control subsystem (TCS) provides passive thermal control of the satellite

system, and also has a backup heater system for atypical operations.

The three primary subsystems, flight computer, EPS, and telecom are designed to work together to resolve

problems resulting from radiation exposure, and unforeseen software faults. First, the EPS subsystem provides

overcurrent protected power to the entire bus. The EPS microcontroller keeps software watch dog timers (SWDT)

on all distribution lines. In an on-orbit situation, the EPS will require that the telecom subsystem and flight

computer check in with EPS periodically to verify their status. If either subsystem does not check in with EPS, the

SWDT expiration results in power cycling the non-reporting device. All other distribution lines have SWDT’s, but

Zenith View

-Y

-X

ZSST-177

C-band

Transponder

Reaction

Wheel

GPS (2x)

Batteries

MD2000C-1

C-band

Transponder

HIP

Camera

OBCS

Telecom

Controller

Telecom

Beacon

EPS

Magtorquer

(3 of 3)

Nadir View

Magtorquers

(2x of 3)

IMU

(2x)

SIP

Camera

Magnetorquer

Control Unit

ADCS

THI

Telecom

Transceivers

Reaction

Wheel Level

Shifter Zenith View

-Y

-X

ZSST-177

C-band

Transponder

Reaction

Wheel

GPS (2x)

Batteries

MD2000C-1

C-band

Transponder

HIP

Camera

OBCS

Telecom

Controller

Telecom

Beacon

EPS

Magtorquer

(3 of 3)

Nadir View

Magtorquers

(2x of 3)

IMU

(2x)

SIP

Camera

Magnetorquer

Control Unit

ADCS

THI

Telecom

Transceivers

Reaction

Wheel Level

Shifter

Figure 14. HawaiʻiSat-1 Internal Configuration

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are turned on and kept alive by the flight computer’s periodic requests. An expired SWDT or power-off request will

result in cutting power from the device.

Secondly, when the flight computer is running normally, the flight computer will monitor and control power to

all devices via EPS. Power to all flight computer controlled devices is done in either a duty cycle method, or

accounted for on an energy credit system. Duty cycle controlled devices are powered on and off in intervals, or

allowed to stay on at all times. An on-demand device (e.g. ADCS magnetic torque rods, or heaters) will have a

finite amount of energy credits assigned to it, and deducted as energy is used by the device. Energy credits

regenerate as time goes on and more power becomes available.

Finally, the telecom subsystem provides an override command service which allows the ground segment to send

specially encoded packets to power cycle the EPS and OBCS. Combined with state-of-health (SOH) beacon updates

from both of those subsystems, the telecom subsystem provides the ground segment with enough information and

control to recover the spacecraft from soft faults in the EPS and OBCS.

ADCS Control

Attitude Status

- Ops Uplink

- Telemetry Downlink

- Beacon Updates

- C&C

- WDT Heartbeat

- SoH Telemetry

- Power Control

3x RS232

RS232 Ethernet

Temperature

Readings

Analog

GPO

Reset Inputs

Power Ethernet

GSE Debug

& Checkout

Interface

Umbilical Power

EPS Latching Relay Control (Safety)

Deployment Detection (Open/Short)

Redundant

GPS

2x RS232

RS232

Beacon

Updates

Power

Onboard Computer Subsystem (OBCS)

TCS

Temperature

Sensors

TCS

Thermostat

Controlled

Thermofoils

Electrical Power

Subsystem

(EPS)

Telecom

Subsystem

Attitude

Determination

and Control

(ADCS)

Lightband

Separation

Connector

Figure 15. HawaiʻiSat-1 Satellite Bus Diagram

Power

- Image Transfer

- Camera Control

- Image Transfer

- Camera Control

Ethernet Ethernet

- Payload Power Control

RS232

Power

CRATEX

XPNDR #1

Onboard Computer Subsystem (OBCS)

CRATEX

XPNDR #2

HSFL Imager

Payload

Primary Camera

(HIP)

HSFL Imager

Payload

Backup Camera

(SIP)

Thermal Imager

Payload

(THI)

- Image Transfer

- Camera Control

Ethernet

Power

Electrical Power

Subsystem

(EPS)

Figure 16. HawaiʻiSat-1 Satellite Bus to Payload Diagram

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Figure 16 shows the bus to payload electrical interfaces. The CRATEX payload has two separate transponders

which will be used in CRATEX operations one at a time. 28V unregulated power is provided to the payloads which

is controlled by the EPS subsystem. The two HSFL imagers, HIP and SIP, receive 12V power from EPS, and

interface with the flight computer to take pictures. Finally, the thermal hyperspectral imager (THI) payload

interfaces to the bus much like the other cameras, but uses 28V unregulated power.

The launch vehicle interface shown in Figure 17 was designed with the launch vehicle engineers to allow a safe

pre-launch checkout. Starting at the top, the satellite’s flight computer and EPS are the only two subsystems which

interface with either the launch vehicle or ground support equipment (GSE). The flight computer’s Ethernet

connection allows the GSE to perform high level diagnostic communication. With this interface, a full satellite

checkout can be performed.

The EPS is designed to receive power from the GSE to charge its battery. There are also latching relays that are

being used to control arming and disarming of the satellite remotely. Finally, a simple logic signal is sent through

the light band separation connector, and looped back at the launch vehicle side. The active logic signal will allow

the EPS to detect deployment events in real time. While on the launch vehicle, the EPS will remain in a safe mode

until a pre-launch checkout is initiated, or the satellite is deployed.

2. On Board Computer Systems (OBCS/ADCS/EPS)

The spacecraft is controllable by real-time uplinked commands, time-delayed command scripts, event-driven

command scripts (e.g. establishing lock with a ground station during a pass), and autonomous onboard flight

software. It has sufficient onboard storage to store at least 24 hours worth of state-of-health (SOH) and payload data.

The spacecraft is capable of storing delayed commands to support the spacecraft and mission for at least 48 hours.

On Board Computer Subsystem (OBCS) consists of two general purpose computers, responsible for the OBCS

and ADCS operations, and a small dedicated processor for EPS functions. The general purpose computers are

400MHz PowerPC MIP 405 based, running on a PC-104+ backplane. The version we have chosen has been flight

tested on both the ISS and the MidStar-1 satellite.4 The dedicated EPS processor is yet to be chosen, but will need to

15" Motorized Lightband

Ethernet to GSE

EPS

Flight

Computer

Ethernet

Switch

Launch Vehicle

Hawai‘iSat-1 Satellite

Ethernet

Deployment

Detection Signal Line

Remote Safing

Control

Shorting

Wire

Battery Charging Power

Launch

Support

GSE

Umbilical

Payload Adapter and Deployer

Motorized

Lightband

Motors

Fireset

PAD/LV Harness

Lightband Motor Control

(Stow, Flight, Deploy)

MLB Control (Deploy Only) Power Safing

Control

Lightband Separation

Connector

MLB Control

(Stow/Flight)

Umbilical

Receptacle

(Pass Thru)

Figure 17. HawaiʻiSat-1 Launch Connections Diagram

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capable of performing sensor sampling and Analog-to-Digital, Digital-to-Analog conversion, and communication

via RS-232. The OBCS Block Diagram is shown in Figure 18.

The functional allocations for the On Board Systems are as follows:

 OBCS MIP405

o Command processing and scheduling

o Wideband telemetry generation

o HSFL imager interfacing and image data compression

o Host payload communications and control

o Temperature monitoring and control

 ADCS MIP405

o Attitude and Position Determination

o Attitude Estimator

o Attitude Control loop

o ADCS sensors interfacing (IMUs, sun sensors, momentum wheel tachometer)

o ADCS actuators interfacing (Magtorquer rods, momentum wheel)

 EPS Controllers

o EPS subsystem control and measurement

o Payload enable / disable

RS232#2

TELECOMM

System

Main CPU

MIP 405

Serial Ethernet

SIP

HIP

Eth#4

Eth#3

Eth#2

EPS

System

THI

RS232#3

SSD

Eth#5

RS422#1

Test Port

RJ45

GPS 2 RS232#4

ADCS

A/D

Eth#1

GPS 1 RS232#5

Temperature

Sensors

Figure 18. OBCS Block Diagram

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3. Electrical Power Subsystem (EPS)

The EPS provides enough power to cover all the spacecraft’s needs during all phases of the mission after launch.

It uses photovoltaic elements (solar cells) assembled into solar arrays to generate power and secondary batteries to

store excess energy until it is required to cover a shortage of power from the solar arrays. The EPS fulfills the power

requirements of the fully operating S/C at all times, including during eclipse periods, for the duration of the mission

(two years) of the spacecraft. The spacecraft generates 28 VDC unregulated current from the solar panels, which is

used to charge the 28 VDC battery. The battery distributes 12 VDC to the HIP and SIP payloads and an unregulated

28 VDC bus to the rest of the subsystems where the local power distribution unit regulates the voltage down to the

required voltage of each subsystem. The EPS Functional Block Diagram is shown in Figure 19.

The eight body-mounted solar array panels each consist of two solar array modules (strings) with 19 solar cells

per module (Figure 20). The solar cells being used are BTJM Triple Junction GaAs solar cells with an efficiency of

28%. The solar arrays generate a peak panel power of ~42 W, with average power over an orbit (with losses) of

~37.7 W. The average power margin over an orbit is ~10.7 W.

The HawaiʻiSat-1 baselined batteries are Panasonic CGR18650 Lithium Ion cells. They are mounted in a 7S4P

(seven series, four parallel) configuration providing 246.9 Wh (8.8 Ah, 25.2 V). These batteries have been flight

proven in the GeneSat-1 and PharmaSat missions.5

The nominal weekday and weekend operation schedule is shown in Figures 21 and 22 respectively. This

determined the loading for the EPS performance shown in Figures 23-26. The nominal S/C attitude is Local Vertical

Local Horizontal (LVLH) hold, which keeps the nadir end (+Z) of the S/C always pointed to geocenter.

EPS uController

Bypass

Out

Main Out

Li-ion Battery

Pack (7S4P)

& Balancer

I-V

Sensor

3.3VDC

Redundant

Regulator

28VDC

Unregulated Bus

I-V Data

HIPSIP

12VDC

Regulated

Switched Bus

28VDC

Unregulated

Switched Bus

I-V

Sensor x9

I-V

Sensor x2

28VDC Bus Enable

12VDC Bus Enable

To ADCS, OBCS,

CRATEX, TCS,

TELECOM, & THI

PV Enable

3.3VDC

I-V

Sensor

Solar Cell

Module 1-a

I-V

Sensor

Solar Cell

Module 1-b

Solar Cell

Module 9-a

Solar Cell

Module 9-b

PV Shunt

Regulator

PV Shunt

Regulator

Figure 19. EPS Block Diagram

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Figure 20. Solar Array Panel Consisting of Two 19-cell Modules

0 200 400 600 800 1000 1200 1400

0.05

0.1

0.15

0.2

0.25

0.3

0.35

Individual Energy Consumpion

Over a Weekday

Time [min]

[WHr]

EPS Bus

Telecom Bus

OBCS Bus

TCS Bus

ADCS PDU Bus

ADCS TCU Bus

CRATEX Bus

HIP/SIP Bus

THI bus

Figure 21. HawaiʻiSat-1 Nominal Weekday Loads

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0 200 400 600 800 1000 1200 1400

0.05

0.1

0.15

0.2

0.25

0.3

0.35

Individual Energy Consumpion

Over a Weekend

Time [min]

[WHr]

EPS Bus

Telecom Bus

OBCS Bus

TCS Bus

ADCS PDU Bus

ADCS TCU Bus

CRATEX Bus

HIP/SIP Bus

THI bus

Figure 22. HawaiʻiSat-1 Nominal Weekend Loads

0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

Total Power Consumption Profile

Over 1 Week, Max: 67.72 [W]

Time [min]

Power [W]

Total

Average Weekday

Average Weekend

Figure 23. Weekly Payload Power Consumption

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010 20 30 40 50 60 70 80 90

Time [min]

Power [W]

Power Generation Profile Over 1 Orbit (LVLH)

Panel 1

Panel 2

Panel 3

Panel 4

Panel 5

Panel 6

Panel 7

Panel 8

Panel 9

Total

Average: 59.21 [W]

Figure 24. Power Generation Profile Over 1 Orbit (LVLH)

020 40 60 80

Time [min]

Power [W]

Solar Array Power Over an Orbit

63% Efficency of DET

Total Power Generated

Average Power Generated 37.30 [W]

Average Power Consumed 26.60 [W]

020 40 60 80

Time [min]

Power [W]

Solar Array Power Over an Orbit

63% Efficency of DET

Total Power Generated

Average Power Generated 37.30 [W]

Average Power Consumed 26.60 [W]

Figure 25. Solar Array Power Over An Orbit

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0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

100

Time [min]

Battery Capacity [%]

Battery Charge Status (246.90 [WHr]) Over 1 Week,

DET Efficiency: 63 [%]

Figure 26. Battery Charge Status Over One Week

4. Telecommunications Subsystem (Telecom)

The preliminary design for the telecom subsystem includes a central telecom controller, and a radio set which is

almost the same as what was flown on the NASA Ames GeneSat-1 and Pharmasat satellites.5 The spacecraft is

capable of two-way communications with the ground in nearly all attitudes. A third radio is currently being sought

for a high speed S-band downlink to transport bulk image data to the ground.

The telecom controller is responsible for management of power, monitoring the temperature of radios, and

listening on all uplink channels for satellite override commands. Override commands include power cycling the

flight computer and the EPS subsystem. The controller will also receive state of health beacon updates from both

the EPS and flight computer. By monitoring both EPS and the flight computer, the telecom subsystem can report

the entire satellite’s status to the ground segment.

The Genesat/Stensat UHF beacon provides amateur frequency beaconing with ASFK. State of health

information will be publicly available for the amateur community to decode and find out more about the satellite. It

is one of our goals to use this beacon for use in educational and community outreach events.

The Microhard MHX-2420 is a newer version of the Genesat MHX 2400 radio, and has a data rate of 115.2kbps.

This is designed in as the primary command and telemetry radio. To keep interfaces simple, RS232 was chosen as

the standard serial signaling between all telecom components.

The ideal characteristics for the third radio include UHF uplink and downlink, and an S-band radio capable of

downlinking data at greater than 768kbps. This would greatly increase the usefulness of all payloads from the

baseline configuration. The UHF link would also allow us to have a redundant command receiver.

The telecom subsystem (Figure 27) was designed with simplicity in mind. The flight computer is the most

capable and efficient for which to develop software. Thus, the uplink/downlink async interfaces are all passed

directly to the flight computer. The flight computer will directly manage the links to the ground.

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TBD UHF

Transceiver

RS232

Transceiver RS232

Transceiver

Microhard ISM-Band

MHX-2420

Transceiver

Genesat UHF

Beacon

Telecom Controller

TBD

Async

Telemetry/

Telecommand

115.2kbps

Beacon Data

1200bps

Async Async

Telemetry/

Telecommand

TBD kbps

Async

GPIO

Monitoring

Async

Telecom Power

Distribution Unit

(EPS Monitored)

Control

Async Async

+5V Distribution

SoH Updates

OBCS Reset Requests

Self Reset Requests

Power Input

From EPS

RS232

RS232RS232

To OBCS To OBCS

To OBCS

Chassis Ground via

EPS only (Star GND)

MMCX TBDSMA-F 1200bps AFSK

T/T

115.2kbps T/T

TBD kbps

T1 T2 T3

Analog

From SE Thermistor T1

From SE Thermistor T2

From SE Thermistor T3

SoH Updates

OBCS Reset Requests

Self Reset Requests

Open Drain GPO

To EPS

EPS Reset RS232

Transceiver

Async

RS232

EPS SoH

Updates

EPS SoH

Updates

To EPS

TEL-CTL1-(TX/

RX) TEL-MHX1-(TX/

RX/CTS/RTS) TEL-TBD-(TX/

RX/CTS/RTS) EPS-DTEL1xxx

TEL-CTL2-(TX/

RX)

EPS-RST-(EPS/

OBCS)

DEVTEL-2 DEVTEL-3 DEVTEL-4

Figure 27. HawaiʻiSat-1 Telecom Subsystem Diagram

5. Structures

The structure ensures all subsystem components are housed within the spacecraft's mass and volumetric footprint,

and adequately protected from physical forces and environment during all phases of the mission, from integration

and testing, through ground transportation, launch, and space flight. The structure also accommodates subsystem

mounting and interconnections.

The design of the satellite structure incorporates a number of primary decks supported by a system of struts (see

Figure 14). The general shape of the structure is an octagon to maximize the power obtained from the solar cells.

The solar cell honeycomb substrate doubles as the side panels for the spacecraft and mount onto the vertical and

horizontal members. The upper solar panels mount onto the zenith surface of the frame. The material to be used in

the structure is Aluminum 7075-T6. A composite structure was considered, but aluminum was selected because of

its reduced cost and risk/complexity. The structure is designed for accessibility, which was part of our criteria for a

standard baseline bus. Helping this are removable, modular solar panels and external electrical/power feedthroughs.

The HawaiʻiSat-1-1 design was also kept mechanism-free for reduced associated risk and complexity.

The design also features a central avionics shelf as shown in Figure 28. This provided modularity, aided assembly

and integration, shortened cable runs, provided additional radiation shielding, and provided stiffening to the S/C

structure.One of the requirements that had to be met in the design of the S/C structure was to not have any object

protrude into the field of view (FOV) of the payload cameras or the sun sensors. This was accomplished as shown in

Figure 29.

The loading used during the design of the structure was 8 g axial acceleration, 7 g lateral acceleration, 2.5 rps

maximum spin rate, and the launch environment and vibration tests specified in MIL-STD-1540E. This resulted in

load limits of 1.25 times the axial and lateral accelerations, a 1.25 factor of safety with respect to yield, and a 1.40

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factor of safety with respect to ultimate. Extensive finite element analyses (FEA) were performed on the HawaiʻiSat-

1-1 structure. The analyses performed included: static (Von Mises Stress), frequency (up to 15th mode), and thermo-

mechanical. An example is shown in Figure 30. However, the analyses were performed using Aluminum 6061-T6

instead of the preferred 7075-T6 alloy. Although this conservative approach built additional margin into the results,

subsequent analyses will be performed with 7075-T6 as the design evolves and matures.

One area in need of attention is the mass of the structure. Utilizing feedback from concurrent engineering

sessions, greater deck thicknesses were used initially with the intention of light-weighting and trimming the deck

thicknesses as designs solidify. A preliminary light-weighting and analysis was in-process at the time of PDR and

yielded mass savings of approximately 10 kg. As this was only an initial effort, it is anticipated that additional mass

savings may be realized.

Figure 30. FEA Showing Axial Displacements

Down to Avionics Deck

GPS (2x)

ADCS

OBCS

Telecom

Controller

EPS

Figure 28. Avionics Shelf

Sun Sensors

THI HIP

SIP

Sun Sensors

THI HIP

SIP

Figure 29. FOV Analysis

Axial (Z) Displacement

Max = 0.17 mm Lateral (X) Displacement

Max = 0.37 mm

Axial (Z) Displacement

Max = 0.17 mm Lateral (X) Displacement

Max = 0.37 mm

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MODE DESCRIPTION

Safe Initial mode in which ADCS will start and for

fault recovery. Only the ADCS processor is

powered on.

Detumble ADCS drives s/c from an arbitrary orientation

and angular velocity to zero angular velocity

and a specific nadir-pointing orientation.

Pointing Spacecraft is commanded to achieve desired

orientation w.r.t. a specific frame along with a

specific angular velocity w.r.t. body frame.

Trajectory Spacecraft is commanded to achieve time-

varying orientation with known angular velocity.

Desaturation Magnetic torque rods generate torque to reduce

angular momentum of RW

Drift Used to create a “quiet” environment for data

collection. Control actuators are off. However,

estimator is operational and retrieving data from

sensors.

Table 3. ADCS Operational Modes

6. Attitude Determination and Control Subsystem (ADCS)

The spacecraft is three-axis stabilized and capable of autonomous, closed-loop inertial pointing with an accuracy

of 3 deg. or better. Attitude measurement accuracy is adequate to determine where the spacecraft is pointing to 1-2

deg. This accuracy is achievable in real-time, in darkness or sunlight, and during all phases of the flight after

deployment.

The ADCS provides the spacecraft with the

capability of maintaining the +z face of the

spacecraft pointing in a nadir (geocentric)

direction with an accuracy of 3. This is the

attitude required for the imaging payloads which

consist of a thermal hyperspectral imager and an

infrared camera. The operational modes of the

ADCS are given in Table 3. Most of the mission

is spent in the pointing (LVLH Hold) mode with

the nadir (+Z-axis) pointing towards geocenter.

However, the spacecraft has the flexibility to

point to any inertial attitude and even to spin

when required. It is also desirable, but not

mandatory, for spacecraft to be capable of

pointing to and tracking a stationary Earth target.

The ADCS software, including algorithms and

settings, are accessible to the ground and can be

updated or replaced from the ground during the

flight if needed.

Attitude data obtained from the spacecraft’s ADCS sensors are autonomously processed aboard the spacecraft to

determine spacecraft orientation. The attitude determination is accomplished using three space-qualified Aero Astro

MSS-01 Sun Sensors, each with 60 full-angle circular FOV providing an accuracy of ~1, and two Microstrain

3DM-GX2 inertial measurement units (IMUs), each of which contains a 3-axis magnetometer, three accelerometers,

and a rate gyroscope.

The ADCS Functional Block Diagram is shown in Figure 31. The commanded control torque vector and the

input vector to the actuators (magnetorquers) are related by: B(R)τa= τc. R is the attitude, τa is the actuator input

vector, τc is the commanded torque vector (from the control law). B(R)=[ Bg B1(R)×Bg B2(R)×Bg B3(R)×Bg] with

three magnetorquers, where Bi(R), i=1,2,3 are the magnetorquer magnetic fields that depend on the attitude, and Bg

is the local geomagnetic field. Whenever one of the magnetorquers is aligned closely with the local geomagnetic

Time and

orbit data

Commanded Trajectory +

Control Law Magnetorquer &

Reaction Wheel

Model

Magnetorquer &

Reaction Wheel Spacecraft in LEO

Estimator Sensor

Data Fusion

Sun Sensors,

IMU,

Magnetometer

Attitude

angular

rate

Disturbance

(atmospheric,

geomagnetic,

solar)

Error in state

τcτaτ

τd

(R*,ω*) (R#,ω#) (R#,ω#)

(R,ω)

Figure 31. ADCS Functional Block Diagram

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Figure 33. TCS Functional Flow Block Diagram

Top

Right

Front

Magnetorquer

IMU

Reaction wheel

Sunsensor

Figure 32. Location of ADCS Components

field, B(R) cannot be inverted to get the τa from the τc. Therefore a small reaction wheel was added for full three-

axis attitude actuation capability.

The ADCS provides attitude control by the use of control actuators, which for HawaiʻiSat-1 are three

magnetorquers and a reaction wheel. The

magnetorquers are space-qualified Vectronic

VMT-35 which each have a magnetic moment

of 35 Am2 @ 100 mA and are controlled by a

Vectronic Torque Control Unit. The reaction

wheel (RW) is a space-qualified Sinclair RW-

0.03-4-ASYNC-2-1-0 which has a nominal

torque greater than 2 mNm and nominal

momentum of 30 mNm-sec @ 5600 rpm. The

layout of the ADCS actuators and components

is shown in Figure 32.

The ADCS control scheme is nonlinear and

continuous, based on feedback of trajectory

tracking errors. Control law τc designed for

global tracking of desired attitude and angular

velocity trajectories.6

The actuator model developed generates

inputs to actuators (RW and magnetorquers) and

the control torque τa applied by the actuators.

This control torque allocation scheme, for a RW along the pitch axis, is given in Reference 7. This scheme is

singular when the local geomagnetic field is perpendicular to the pitch (RW) axis.

The ADCS estimation and filtering scheme is nonlinear and discrete, based on deterministic ellipsoidal bounds

on measurement errors and dynamic flow uncertainty. Measurements are: direction vectors (for attitude) and angular

velocity vector, all in spacecraft body frame. Estimation scheme for an orbiting satellite without control torques

given in Reference 8. Propagation of attitude and angular velocity between measurements carried out using a Lie

Group Variational Integrator.

The pointing (LVLH Hold) attitude mode requires a pitch rate of 0.063 º/s. Simulations of a detumbling

maneuver that concludes with stabilization about a nadir-pointing attitude carried out by ADCS shows satisfactory

de-tumbling from an initial attitude error angle of 90º and initial angular velocity of 0.098 rad/s within 1/20 of an

orbit at 550 km altitude. The detumble attitude mode is capable of achieving a desired orientation to a nadir

orientation in the LVLH frame with a total required torque magnitude < 0.3 milli-N-m and reaction momentum less

than 3.5 milli-N-m-s.

7. Thermal Control Subsystem (TCS)

The TCS is designed to ensure all spacecraft

subsystem components are thermally controlled for

operation, both in sunlight and eclipse periods. The

TCS will also provide SOH temperature from all

critical points of the spacecraft to the OBCS. Due to

the fairly benign thermal environment of low Earth

orbit, passive thermal control was determined to be

sufficient. The internal power dissipation required to

be handled by the TCS is from a minimum of 2.6 W to

a maximum of 96.2 W. The TCS hardware consists of

multi-layer insulation (MLI) blankets on the interior

S/C panels, appropriate surface finishing (e.g., paints),

three thermostats with redundancy, heaters for critical

components and temperature sensors (RTDs). The

TCS Functional Flow Block Diagram is shown in

Figure 33.

The A/D board that will collect the thermal data is capable of reading 32 single-ended inputs and is compatible

with resistive thermal device sensors (RTDs) making is possible to have a significant amount of thermal information

from the various subsystems in the spacecraft. This information is passed to the OBCS for post processing and

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Figure 35. Incident Heat Flux for One Orbit

storage. The OBCS is capable of deciding if some power line for the heaters must be turned on or off in case of any

failure from the thermostats or the EPS. The hardware schematic is shown in Figure 34.

The incident heat flux is shown in

Figure 35 for one orbit. This shows

clearly in the analysis where the hot

and cold faces for the satellite will

be.

Detailed simulations were

performed of the spacecraft including

all sub-systems and payload using

Thermal Desktop using finite

differences. Simulations were done

for a nominal hot case and a cold

case. Trades studies for the thermal

management changed the layout of

internal components. Figures 36-38

shows the runs for the ADCS

subsystem (cold, nominal and hot

cases respectively).

Figure 34. TCS Hardware Schematic Diagram

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Figure 38. Simulated Temperature of the ADCS – Hot Case

Figure 37. Simulated Temperature of the ADCS –Nominal Case

Figure 36. Simulated Temperature of the ADCS – Cold Case

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8. Flight Software(FSW)

The flight software autonomously monitors and maintains the spacecraft’s state of health, controls the operation

of the spacecraft, monitors the SOH of the payloads, performs the calculations for the ADCS, and can activate or

deactivate the payloads. The FSW will always be recoverable from loss of power or function without ground

support. It can also be reloaded or modified from the ground without endangering the spacecraft SOH or mission.

The FSW architecture is shown in Figure 39.

Tasks that require persistency are handled by continuously running processes. All other tasks are handled by

singly run programs. FSW modes are handled as a state machine amongst the persistent processes. All commands

are handled as independent programs. These programs either request changes in state from the processes, or directly

perform simple functions themselves. An executive process oversees the launching of all sub processes and

programs. It is the primary path of communications in to the software.

The FSW is spread across the processors on board the spacecraft. The OBCS portion is responsible for command

and control, communications, spacecraft health, and instrument control. The ADCS portion performs the navigation

calculations for the spacecraft to generate an orbit and position that is used by other subsystems. The current

position of the spacecraft will data obtained from one of two NovaTel GPS units, backed up by regularly uploaded

Two Line Element (TLE) sets. Thus the S/C will be able to determine its position even if it is out of touch with the

ground for extended periods. On both the OBCS and ADCS, the FSW operating system is a standard Linux running

a 2.6 kernel. Finally, more autonomous systems, like the EPS, will run an embedded single string of execution,

dedicated to their particular task.

IV. Payloads

This section describes the payloads being carried by the HawaiʻiSat-1 spacecraft.

A. C-Band Radar Transponder Experiment (CRATEX)

The CRATEX experiment payload is designed to flight test a replacement to the RADCAL payload currently

operating in a degraded fashion on the RADCAL satellite, which was launched in 1993 from Vandenberg Air Force

Base (VAFB) with an expected lifetime of two years. HawaiʻiSat-1 with the CRATEX payload will test some new

technology that can be used to support ground based Range Instrumentation’s performance monitoring functions.

Ground based Range Instrumentation users are minimally comprised of Tri-services and NASA Test Ranges.

Although the optimal satellite orbit is 850 Km altitude, with an inclination of 65° or higher, the HawaiʻiSat-1 orbital

altitude of 550 km in a Sun-Synchronous (inclination ~98°) is considered by VAFB to be satisfactory for testing

purposes. A C-Band transponder is used to calibrate the ground radars. An orbit determination system is required

that achieves a minimum of 5 meter accuracy. The orbit may be determined on-board, on the ground or through a

FLIGHT CPU

MIP 405

Executive

SOH Telecomm Payloads

UDP Sockets

POSIX IPC

and Signals

System

Daemons

Commands Data

Figure 39. Spacecraft Flight Software Architecture

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Figure 40. SST 177C C-Band Transponder

Figure 41. MD2000C C-Band Transponder

Figure 42. Prosilica GS2450C

combination of space and ground calculations, but must be available to users within 48 hours. The primary method

used on LEO-1 for orbit determination is a NovaTel GPS receiver.

VAFB is supplying two C-band transponders – the

space-qualified SST-177C (currently being flown on

DMSP 15 satellite) and the transponder to be flight-

tested in space, the MD2000C (Figures 40 and 41

respectively). Two AntDevCorp Quadrifilar Helix C-

Band antennas are being used. Both transponders

receive ground C-Band radar signal of 5765 MHz and

respond coherently with 5765 MHz on the SST 177C or

at 5690 MHz from the MD2000C. This allows

monitoring of the tracking performance of C-Band

radars, while testing the performance of the MD2000C

in comparison to the older technology of the SST 177C.

Radar data (TRAE) is compared with a known satellite

position (precise ephemeris) obtained from processed

GPS data. A coherent transponder is necessary for

radars equipped with range rate subsystem. The satellite

provides a dynamic target for testing the total radar

system. C-Band transponders are normally off and must

be commanded on just prior to requested use and off

again immediately after each requested use. Track

period is software limited to a maximum of five ON

segments of no more than 20 minutes in a 24 hour

period.

Ephemeris must be determined to better than five

meters accuracy. Once a day, the data are transferred to

Western Range VAFB for initial editing and processing.

The edited data are then forwarded to the National

Geospatial-intelligence Agency (NGA) for final

processing and posting of resultant ephemeris on

RPMWEB site for user community access.

B. HSFL Imagers: HSFL Imager Payload (HIP) and Separation Imager Payload (SIP)

The HSFL Imagers provide imagery that fulfills a secondary mission objective and will be used by HSFL for

public outreach and to inspire interest in science education. The final product of this experiment includes (1) a color

image showing the Hawaiʻian Islands with little or no cloud cover. This can be done with either a single image or a

mosaic of images primarily using the HIP camera; and (2) color images to document the separation of HawaiʻiSat-1

from the PAD during orbital deployment. This latter activity is the primary purpose of the SIP camera, but it would

subsequently be a backup camera for the HIP camera. Both

are nadir-pointing (in normal LVLH Hold mode).

Both Imagers are Prosilica GS2450C color 2448 x 2050 (5

MP) CCD image sensors with a narrow FOV (NFOV) lens

placed on nadir-pointing side of spacecraft (see Figure 42).

Its frame rate is up to 15 FPS. The lens selected for HIP is an

Edmund Optics NT57-680 12-36 mm focal length lens with a

FOV of 15. Its focus range is 200 mm to infinity. The Lens

for the SIP Imager is a Pyramid Imaging C30405KP lens

with a 4.8 mm focal length. The C30405KP is the

replacement for the older C30405TH part which has flown on

the Rocket Observations of Pulsating Aurora (ROPA)

sounding rocket, on the De-spun Rocket Borne Imager 2.9

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Figure 43. THI Configuration

Figure 44. THI Optics Method

C. Thermal Hyperspectral Imager (THI)

The Thermal Hyperspectral Imager (THI) Payload is a hosted scientific payload from Hawaiʻi Institute of

Geophysics and Planetology (HIGP) at the University of Hawaiʻi at Manoa. This payload consists of an imager and

computer inside a pressure vessel, an interferometry cube, and a calibration system. The configuration of the THI as

it will be mounted in the satellite is shown in Figure 43. HSFL shall provide THI will power and data interfaces.

The integration of the THI payload onto HawaiʻiSat-1 is the final segment of a research project funded by a

NASA EPSCoR grant to develop a hyperspectral imager that is both lighter and more power efficient than previous

imaging systems so as to be easily integrated into a small satellite for earth observation from orbit. The ability to

acquire data from orbit will allow a more complete and uniform data set of Earth’s spectral reflectance and

emittance. The data acquired will provide valuable information pertaining to environmental issues such as coral reef

health, pollution and volcanic activity, all issues which directly affect the state of Hawaiʻi.

The instrument will be capable of acquiring

calibrated radiance data at visible and near

infrared (VNIR, ~0.4-1.0 μm) and thermal

infrared (TIR, 8-14 μm) wavelengths. The

Instrument being used will be converted from

instrumentation used on airborne imaging

systems that have been built and flown by

scientists at HIGP. The interferometry cube for

collecting the TIR data consists of three mirrors

on three contiguous sides of the cube and a

ZnSe beam splitter, BS, which is allowed to

rotate (as seen in Figure 44). The rotation of the

beam splitter imposes a varying optical path

across the detector array which creates the

interferogram. The VNIR instrument has been

built by a local Hawaiʻi business based on

systems used by the Civil Air Patrol.

Calibration is required before every data

acquisition at the least and may also be

completed after the acquisitions. The calibration

system consisted of a hot paddle and a cold

paddle that can be swung into the imager’s field

of view. The temperature of the paddles will be

known so that the images acquired of the

paddles will provide a key to the images

acquired of the Earth after that calibration.

HSFL will be testing the pressure vessel, the

electrical, mechanical, and data interfaces. The

THI software is being developed in close

calibration with the HawaiʻiSat-1 software team

so few interface issues are expected.

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V. Operations

This section describes the operations concept for the HawaiʻiSat-1 mission including ground stations.

A. Operations Concept and Phases

The HawaiʻiSat-1 mission is divided into four major segments during the mission Phases E and F.

1. Pre-Launch and Launch Operations

This covers the activities required after the HawaiʻiSat-1 has been completed and put into storage ready for

launch, and through launch up to the deployment of the satellite in orbit. The specific tasks involved are:

 Storage and removal from storage of the satellite.

 Checkout of the satellite to verify that it is fully functional and ready for the mission.

 Integration to the PAD (may possibly be done at the launch site).

 Shipping of satellite (and possibly PAD) to the launch site.

 Integration with the launch vehicle within its fairing.

 Pre-launch checkout.

 Launch and ascent to orbit.

 Deployment from PAD and launch vehicle.

The HawaiʻiSat-1 satellite will be launched partially activated ready to image the deployment from the PAD.

There is a definite sequence of events during deployment from the launch vehicle, including imaging the separation

with the SIP camera. The attitude of the satellite will be determined and the attitude stabilized as the satellite

transitions into a nominal orbit mode.

2. Engineering Evaluation & Checkout (EE&C)

This is the checkout period of the HawaiʻiSat-1 after orbit insertion and is expected to last about a month (or as

long as necessary to complete the checkout). This phase concludes with its commissioning when it has achieved

initial operational capability (IOC). During the EE&C the following tasks are performed:

 Testing of spacecraft bus subsystems and modes.

 Testing of ground segment and data flow.

 Testing of CRATEX payload:

- Both C-band transponders and antennas tested

- Orbit determination method from onboard GPS data tested for 5-m accuracy

- Test interfaces with USAF for tasking and delivery of orbit data

 Testing of THI payload:

- Determine on-orbit calibration of the THI

- Take images of various sizes and with various image settings – downlink images to ground

- Test onboard compression and compare with uncompressed images

 Testing of HSFL Imagers:

- Take images using real-time and delayed commands using both cameras with storage and downlink

of images.

- Take images with various camera settings, such as integration time.

3. Nominal Operations

Nominal operations are the spacecraft operations that occur after IOC and until end of mission (Primary and

Extended missions). Most of this phase of the mission the spacecraft will be spent in what is called nominal mode.

The nominal mode for the HawaiʻiSat-1 spacecraft is to support the CRATEX and THI payloads. This mode has the

HawaiʻiSat-1 in LVLH hold (+X or –X forward) attitude (see Figure 45). The operational modes within nominal

operations are:

 Nominal Mode – CRATEX Operations

- Each ground radar requires calibration at least once per week (using Monday to Friday schedule).

- A weekly schedule request is provided by VAFB (30 Space Wing) >3 days in advance. HSFL

integrates requests with the overall HawaiʻiSat-1 schedule. The final CRATEX contact schedule is

provided to 30 SW by the HMOC at least two days in advance for nominal commanding.

- A C-band transponder is turned on and off by time-delayed script before and after each calibration

contact (script nominally uploaded >1 day in advance). Only one transponder and one antenna are

active at a time. Transponders and antennas 1 and 2 are used in alternating months. Transponder 1

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(SST 177C) is the reference transponder and Transponder 2 (MD2000C) is the experimental

transponder.

- There are a maximum of five calibration passes per day (each of 20 minutes maximum duration

limited by watchdog timer).

- The target ground radar activates the C-band transponder and executes a calibration contact.

- HSFL sends GPS position data to VAFB.

 Nominal Mode – THI Imaging.

The THI camera uses a “push-broom” technique to create an image along the nadir ground track of

constant width but variable specified length. For this experimental version of THI, no off-track pointing

is required, although would be possible using the ADCS capability of the HawaiʻiSat-1. Since THI is a

thermal imager, it will normally be operated during darkness, although daylight imaging is possible.

 Nominal Mode - Earth Imaging.

In Nominal Mode either HSFL camera (i.e., HIP or SIP) can take images of the Earth. The HIP camera

has a comparatively narrow FOV and will be the primary camera used during the mission for Earth

imaging. The similar SIP camera is primarily used to image the separation of the HawaiʻiSat-1 from the

launch vehicle, but can also be used as a backup for the HIP imager for obtaining images of the Earth,

although with lower resolution and a wider-angle lens. However, the HSFL imagers are not restricted to

nadir imaging. The ADCS is capable of pointing the spacecraft in any direction or even to track a

ground target, which provides much greater flexibility to the imaging targets of the HSFL cameras.

Other than looking at off-track targets, the cameras would also be capable of pointing to celestial targets,

such as calibration starts or the Moon. Images can be taken by time-delayed script or real-time command

on a non-interference basis. They will be stored onboard and then downlinked when convenient.

Having determined the various operational modes to fulfill the objectives of the payloads, a weekly schedule is

formed that ensures that there are no conflicts between the spacecraft bus and the various payloads. A nominal

weekly schedule showing opportunities for the operations of the various payloads and modes is shown in Figure 46.

LH LH

Sun Vector

8 6

NOMINAL MODE

Attitude: LVLH Hold, +x (or –x)

+x SA: 1,8,7,6,5 sunlit periods

2,3,4 dark always

-x SA: 5,4,3,2,1 sunlit periods

6,7,8 dark always

Imagers: Nadir

LV = Local Vertical

LH = Local Horizontal

LH LH

Sun Vector

8 6

NOMINAL MODE

Attitude: LVLH Hold, +x (or –x)

+x SA: 1,8,7,6,5 sunlit periods

2,3,4 dark always

-x SA: 5,4,3,2,1 sunlit periods

6,7,8 dark always

Imagers: Nadir

LV = Local Vertical

LH = Local Horizontal

Figure 45. Nominal Mode

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4. Mission Termination/Disposal

New missions launched into LEO must have a means to dispose of the satellite within 25 years of the mission

termination in order to remove it as potential space debris. This can be done either using a system to change the

orbital velocity of the satellite (e.g., using propulsive or drag devices), or by natural orbital decay. For HawaiʻiSat-1,

which has no propulsion subsystem, the latter method was selected. The current analysis shows that with an initial

altitude of 550 km, that the satellite will reenter within 25 years of the nominal mission end, even with uncertainties

in the actual length due to factors such as solar activity and orbit insertion dispersions.

At the end of mission, the spacecraft payloads and transmitter will be turned off. There is no further action

required before reentry into the atmosphere.

B. Operational Modes

A number of operational modes have been identified for the HawaiʻiSat-1 spacecraft to accomplish its mission.

These operational modes are identified and described in Table 4. There is a close correlation between these modes

and the ADCS modes that were identified in Table 3. The specific operational modes used on-orbit and the transition

between them is shown in Figure 47.

Instrument MONDAY

TUESDAY

WEDNESDAY THURSDAY FRIDAY

SATURDAY

SUNDAY

CRATEX

THI

HSFL Imagers

Figure 46. Nominal Payloads Weekly Schedule Showing Opportunities

Table 4. Spacecraft Operational Modes

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C. Operations Staffing

The basic staffing to support operations of the HawaiʻiSat-1 mission after launch is shown in Table 5. The basic

operations philosophy of the HawaiʻiSat-1 mission is to use the spacecraft engineers as operations personnel during

the EE&C period of the first month until IOC. During nominal operations, the engineers will be working other

projects and will be used to support operations as needed (e.g., anomaly resolution) or to just a very low level. Most

of the operations will be performed by the spacecraft controllers, which will be mainly students. This also provides

an educational benefit to the project, to familiarize students with the operation and workings of an operational

satellite. The spacecraft is being designed for mostly autonomous operation, including automatic contacts with the

ground where command files are uplinked, and data files are downlinked with human intervention.

Once the mission has achieved a certain level of maturity in which there is a high level of confidence that the

automated systems will work, then the mission can enter the mature operations, which requires minimal staffing by

operations personnel. Only a single 8-hour shift for five days a week is anticipated for nominal and mature

operations, while the EE&C period will require operations personnel for seven days a week.

Figure 47. On-orbit Operational Modes and Transitions

EE&C Nominal Mature Shift

Position FTE FTE FTE h/d Comment

Operations Manager 1 0.5 0.1 8/5

Spacecraft Controller 2 0.9 0.1 8/5 8/7 for EE&CO

Ground Network Controller 1 0.1 0.1 8/5

Scheduler 1 0.5 0.5 8/5

Mission Planner 1 0.5 0.5 8/5

Data Manager 1 0.5 0.1 8/5

S/C Analyst 4 0.5 0.1 8/5 8/7 for EE&CO

Software Analyst 2 0.5 0.1 8/5 8/7 for EE&CO

Orbit Analyst 1 0.5 0.1 8/5

TOTAL 14 4.5 1.7

Table 5. HawaiʻiSat-1 Operations Staffing

American Institute of Aeronautics and Astronautics

D. Operations Process

The mission operations process developed for HawaiʻiSat-1shown in Figure 48 includes the following four sub-

processes:

1. Planning Process

This takes the Mission Operations Plan, and the Flight Rules & Mission Requirements (all developed before

launch), with the current status of the spacecraft, the ground network, and any payload customer requests

(such as CRATEX contact requests received from 30 Space Wing at VAFB) and with the aid of the current

orbit ephemeris and visibility windows, determines a master schedule, and generates timelines for orbits as

necessary. From the timeline a command script is automatically generated, and if necessary a set of pointing

quaternions. The command script and quaternions are then verified in the Operational Testbed

(OTB)/simulator, and if successful, passed along to the Execution Process for implementation.

2. Execution Process

This handles the real-time operations of the ground network (ground stations and network) and the

interactions with the HawaiʻiSat-1 spacecraft. Contacts with the spacecraft are monitored, during which the

command script and flat files are uplinked, while the SOH (both real-time and archived) and payload data are

downlinked to the ground station and transferred to the HMOC. All contacts and events are logged, both

automatically and manually by the controllers (if present).

3. Data Management Process

This takes the downlinked SOH and payload data from the ground station, then archives, processes, and

distributes the data.

4. Analysis Process

The processed data are analyzed and trended over time to determine the health of the spacecraft and the

success in accomplishing mission objectives. If anomalies are detected, then a spacecraft or payload engineer

is alerted to initiate an anomaly resolution process.

DATA MANAGEMENT

PROCESS

Contact Plans

Command Loads/Scripts R/T

Commands

(GN & S/C)

All SOH Telemetry

GN or S/C Track Data

R/T FLIGHT OPERATIONS

PROCESS

Support Schedule

Payload Data

Products & S/C Constraints

ANALYSIS PROCESS

PLANNING

PROCESS

Commands, Tasking, Constraints

S/C SOH Data

Anomalies &

Eng. Data

Anomalies

Resolution

& Reports

Reports, Iterative Schedule

GN Telemetry & Track Data

R/T SOH Data

(S/C & GN)

All Level 0 &

SOH Data Data

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